1. Technical Field
This invention relates generally to a method of compensating for pointing of an orbiting satellite, and more particularly, to a method for compensating for the effects of solar pressure on the pointing direction of an orbiting satellite as the satellite goes into and out of eclipse by the sun.
2. Discussion
A geosynchronous earth orbit, as is known in the art, is the orbit about the earth in which a satellite or spacecraft will remain fixed above a specific location on the earth. This orbit is at a distance of approximately 22,400 miles above the earth. In this orbit, a beam, such as a communications beam, from the satellite can be maintained over a desirable area on the earth, such as a particular country, thus establishing an area which will receive the beam. To remain in a geosynchronous orbit it is necessary that the satellite be in an orbit substantially within the equatorial plane of the earth at the desirable distance, and that the satellite's attitude be oriented perpendicular to this plane. Any deviation or disturbance which causes the satellite to direct its antenna away from a boresight location on the earth tends to effect the coverage area of the beam, and thus, produces undesirable results. Many different forces are in effect on the satellite which tend to alter the satellite's antenna pointing direction.
As a first order method for countering the effects of the different forces acting on the satellite, it is known to stabilize the satellite's attitude by providing an angular bias momentum which resists changes in the satellite's orientation due to external forces transverse to the bias momentum axis. Satellites using this technique are generally referred to as "momentum bias" satellites. Angular momentum bias is usually provided by a number of momentum or reaction wheels which spin at least part of the satellite. The bias axis set by the spin of the momentum wheels is generally perpendicular to the direction of the orbit of the satellite. Although the bias momentum resists changes in the satellite's orientation in directions transverse to the bias momentum axis, it is still necessary to provide control for correcting variations in the satellite's orientation along the bias axis. Different methods of controlling the satellite's attitude, such as feedback loops, are known in the art.
For most bias momentum satellites, the satellite payload, i.e., the part of the satellite carrying at least the antenna, is oriented differently than the momentum wheel. It is therefore necessary to provide means for correcting the orientation of the payload with respect to the orientation of the momentum attitude. Typically, the satellite's payload is defined in three axes referred to as the yaw, roll and pitch axes. If the satellite is in a geosynchronous orbit, the yaw axis is generally directed from the satellite to the center of the earth, the pitch axis is generally directed normal to the plane of the orbit of the satellite and the roll axis is generally perpendicular to the yaw and pitch axes, in a direction of travel of the satellite as is well known in the art.
One particular disturbance torque which affects the pointing of a geosynchronous satellite is the disturbance torque which can be equated to solar pressure caused by particles from the sun hitting the satellite. A satellite in a geosynchronous orbit will be eclipsed from the sun one time per revolution of the earth, and thus, the solar pressure on the satellite will go to zero during the eclipse. Typically, the attitude control mechanisms in use on the satellite will compensate for the effect of solar pressure. However, on a momentum bias satellite without direct yaw sensing, disturbance torques are typically estimated with relatively long time constants. These estimated disturbance torques are used by the satellite's attitude control system to create equal and opposite control torques to eliminate or reduce satellite attitude transients. Compared with these time constants, the solar pressure disturbance torque undergoes a rapid transient during eclipses, both when going from full solar pressure disturbance to zero disturbance, and then returning to full disturbance from zero, generally within a matter of minutes. This change is too fast for the disturbance torque sensing to track, and thus leads to potentially large satellite attitude disturbances, especially about the yaw axis.
What is needed then is a method to calculate the changes in solar disturbance torques caused by eclipses in order to compensate for this effect on the pointing of the satellite. It is therefore an object of the present invention to provide such a method.